The present invention relates to thermal barrier coatings, a method, and an apparatus for determination of past-service conditions of coatings and parts and remaining life thereof. In particular, the present invention relates to such a method and an apparatus by a non-destructive optical determination of a particular crystalline phase in a thermal barrier coating.
The constant demand for increased operating temperature in gas turbine engines has necessitated the development of ceramic coating materials that can insulate the turbine components such as turbine blades and vanes from the heat contained in the gas discharged from the combustion chamber for extending the life of such components. These ceramic coatings are known in the art as thermal barrier coatings.
A thermal barrier coating typically comprises at least a layer of a refractory or thermally insulating material such as yttria-stabilized zirconia (or “YSZ”) which is zirconia stabilized with, for example, about 6-8 percent by weight of yttria. The refractory material would generally be selected to have a low thermal conductivity such as about 1-3 W/(m)(K), thereby reducing heat transfer to and the temperature experienced by the turbine engine component. The coating may be applied by one of known deposition techniques such as the thermal or plasma spray process or the physical vapor deposition process. A typical thermal barrier coating is a multilayer system comprising three layers. A first so-called bondcoat is applied to the surface of the superalloy of the turbine component. This bondcoat typically comprises a MCrAlY alloy wherein M is nickel, or cobalt, or PtNiAl alloys. The purpose of the bondcoat is to provide a layer which adheres well to the underlying alloy, which provides protection against oxidation of the alloy, and which provides a good base for further coatings. A second intermediate layer or interlayer is applied on the bondcoat. A suitable material for this interlayer is Al2O3. This material can be formed by oxidizing the surface of the bondcoat to form an oxide layer. The interlayer provides improved adhesion for the final thermal insulating YSZ coating and is not included for a thermal barrier property.
Despite great care taken during manufacture to ensure good adhesion of the thermal barrier coating to the underlying material of the turbine component, thermal cycling during use of such a component eventually leads to spalling of the coating. In addition, erosion of the thermal barrier coating is inevitable over an extended period of use. Such a spalling or erosion would eventually expose the underlying alloy to extreme temperatures that would lead to failure of the component. Therefore, thermal barrier coatings need be inspected frequently for any sign of deterioration. Such an inspection often requires taking the engine component out of service and is time-consuming. A common inspection technique is the visual inspection of the presence or absence of coating. While that method determines when a spall has occurred, it is unable to determine either the degree of deterioration in an intact coating. A method for determining the past-service conditions and remaining life of thermal barrier coatings would be welcome in the art.
Similarly, it is desirable to monitor the condition of the turbine components themselves. In the prior art, it is usual for a destructive evaluation to be performed at each inspection interval for critical components in the hot gas path. In that case, one part is destroyed to produce sections for metallographical examination. The condition of the coatings and base materials are determined from metallographical inspection, and a decision to repair or replace the remaining parts is made from that information.
Better knowledge of the past-service conditions experienced by the turbine components would allow the determination of the remaining life of a part without destructive evaluation. Currently there are few in-situ measurements of hot gas path parts temperatures available. Some physical changes in the phases and structures of the materials of thermal barrier coatings and components occur with exposure to high temperatures. Inspection for changes in phase content is one way to determine past-service conditions.
However, traditional methods of inspection, such as X-ray diffraction and neutron diffraction, require destructive testing and specialized equipment. They are not conducive to being deployed at the site of a gas turbine. In addition, such destructive testing methods necessarily extrapolate the result obtained for one part to the condition of other similarly used parts and, thus, may not provide a true and accurate condition of those parts.
European patent application EP 0863396 A2 discloses a non-destructive measurement method for residual stress proximate an interlayer in a multilayer thermal barrier coating system. This method focuses on detecting compressive stresses that accumulate at the boundary between the interlayer and the outermost thermal barrier coating by detecting the shift in frequency of light emitted by fluorescing chromium ions in the alumina interlayer. However, significant stresses at that boundary may not appear until after the outermost barrier layer has seriously deteriorated. Furthermore, the stresses at the boundary are not useful indicators of the past-service conditions of the component itself. Therefore, such a method is not very useful in timely forewarning a need for repairing or replacing the engine component.
Therefore, there is a continued need to provide a simple non-destructive method for determining the past-service condition of a thermal barrier coating of a component used at high temperature in a turbine engine. It is also very desirable to provide a method by which the remaining useful life of the underlying component may be determined or estimated. Furthermore, it is also very desirable to provide such a method so that maintenance of turbine engine components may be performed only on an as-needed basis rather than on a fixed schedule.